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National Advisory Committee for Aeronautics 



NO. 92 



R 



esearc 



h Abstracts 



NOVEMBER 15, 1955 



CURRENT NACA REPORTS 

NACA Kept. 1175 

EFFECT OF VARIABLE VISCOSITY AND THERMAL 
CONDUCTIVITY ON HIGH-SPEED SLIP FLOW BE- 
TWEEN CONCENTRIC CYLINDERS. T. C. Lin and 
R. E. Street, University of Washington. 1954. ii, 
36p. dia^rs. (NACA Rept. 1175. Formerly 
TN 2895) 

The dillerential equations ol slip flow, including the 
Burnett terms, were first solved by Schamberg as- 
suming that the coefficients of viscosity and heat con- 
duction of the gas were constants. The problem is 
solved herein for variable coefficients of viscosity 
and thernjal conductivity by applying a transforma- 
tion leading to an iteration method. The method, 
starting with the solution for constant coefficients, 
enables one to approximate the solution for variable 
coefficients very closely after one or two steps. 
Satisfactory results are shown to follow from 
Schaniberg's solution by using his values of the 
constant coefficients multiplied by a constant 
factor n, leading to what are denoted as the effec- 
tive coefficients of viscosity and thermal conduc- 
tivity. 



NACA Rept. 1189 

THEORETICAL AND EXPERIMENTAL ANALYSIS 
OF LOW-DRAG SUPERSONIC INLETS HAVING A 
CIRCULAR CROSS SECTION AND A CENTRAL 
BODY AT MACH NUMBERS OF 3. 30, 2. 75, AND 
2.45. Antonio Ferri and Louis M. Nucci. 1954. 
ii, 37p. diagrs., photos. (NACA Rept. 1189. 
Supersedes RM L8H13) 

Contains theoretical and experimental analysis of 
circular inlets having a central body at Mach num- 
bers of 3.30, 2.75, and 2.45. The inlets have been 
designed in order to have low drag and high pressure 
recovery. The pressure recoveries obtained are of 
the same order of magnitude as those previously 
obtained by inlets having very large external drag. 



NACA Rept. 1198 

A THEORETICAL STUDY OF THE EFFECT OF 
FORWARD SPEED ON THE FREE-SPACE SOUND- 
PRESSURE FIELD AROUND PROPELLERS. I. E. 
Garrick and Charles E. Watkins. 1954. ii, 16p. 
diagrs., tab. (NACA Rept. 1198. Supersedes 
TN 3018) 

The sound-pressure field due to thrust and torque of 
a propeller in flight at uniform subsonic speed is 




analyzed by use of a distribution of acoustic doublets 
located at the propeller disk. The basic element 
used to synthesize the field is the pressure field of a 
concentrated force moving uniformly at subsonic 
speeds, for which an expression generalizing one of 
Lamb's for the fixed concentrated force is given. 
The results can be regarded as an extension of the 
work for the static propeller given by Gutin (NACA 
TM 1195) for the far pressure field and given by 
Hubbard and Regier (NACA Rep. 996) for the field 
near the tips of the rotating propeller. The extendfed 
formulas are used to calculate the sound field for a, 
specific two-blade propeller operating at constant \ 
torque for various forward-speed Maqh,fl\ynbers. 



NACA Rept. 1209| 

DEVELOPMENT OF TURBULENCE-MEASURING 
EQUIPMENT. Leslie S. G. Kovasznay, John 
Hopkins University. 1954. ii, BOp. diagr§.. 
photos., tab. (NACA Rept. 12091 Supersedeg,. 
TN 2839) '"S^ 

Hot-wire turbulence-measuring equipment has been 
developed to meet the more-stringent requirements 
involved in the measurement of fluctuations' in flow 
parameters at supersonic velocities. The higher 
mean speed necessitates the resolution of higher fre- 
quency components than at low speed, and the rela- 
tively loA' turbulence level present at supersonic 
speed makes necessary an improved noise level for 
the equipment. The equipment covers the frequency 
range from 2 to 70,000 cycles per second. The 
equipment is adaptable to all-purpose turbulence 
vvork with improved utility and accuracy over that of 
older types of equipment. Sample measurements 
are given to demonstrate the performance. 



NACA RM E55I16 

SPARK IGNITION OF FLOWING GASES. V - AP- 
PLICATION OF FUEL-AIR-RATIO AND INITL^L- 
TEMPERATURE DATA TO IGNITION THEORY. 
Clyde C. Swett, Jr. November 1955. 19p. diagrs. 
(NACA RM E55I16) 

Data showing the effect of fuel-air ratio and initial 
temperature on spark-ignition energy are presented 
and applied to a previously developed theory of igni- 
tion. The initial-temperature data are consistent 
with the theory; fuel-air-ratio data are only par- 
tially consistent. Probable reasons for the discrep- 
ancy are discussed. 



•AVAILABLE ON LOAN ONLY. 

ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1512 H ST., NW,, 

THE REPORT TITLE AND AUTHOR. 



WASHINGTON 25, D C , CITING CODE NUMBER ABOVE EACH TITLE, 






NACA TM 1388 

GENERAL SOLUTIONS OF OPTIMUM PROBLEMS 
IN NONSTATIONARY FUGHT. (Soluzioni General! 
di ProWemi dt Ottimo in Volo Non-Stazionario). 
Angelo Miele. October 1955. 25p. dlagrs., tab. 
(NACA TM 1388. Trans, from L'Aerotecnica, v.32, 
no.3, 1952, p. 135-142) 

A general method concerning optimum problems in 
nonstationary flight is developed and discussed. 
Various conditions of flight in a vertical plane 
(climb with minimum time, climb with minimum fuel 
consumption, steepest climb, descending and gliding 
flight with maximum time of space) are studied; 
the corresponding best techniques of flight, that is, 
the optimum speed-height relationships, are deter- 
mined. 



NACA TN 3415 

A UNIVERSAL COLUMN FORMULA FOR LOAD AT 
WHICH YIELDING STARTS. L. H. Donnell and V. C. 
Tsien, Illinois Institute of Technology. October 
1955. 48p. diagrs., photos., tab. (NACA TN 3415) 

An analysis is presented of the load at which yielding 
first occurs in actual columns, taking adequately into 
account all the factors which have an important ef- 
fect upon this load. The results are expressed as a 
formula or chart applicable to all cases. 



NACA TN 3476 

CALCULATED SPANWISE LIFT DISTRIBUTIONS 
AND AERODYNAMIC INFLUENCE COEFFICIENTS 
FOR SWEPT WINGS IN SUBSONIC FLOW. 
Franklin W. Diederich and Martin Zlotnick. 
October 1955. 173p. diagrs., tabs. (NACA 
TN 3476) 

Spanwise lift distributions have been calculated for 
61 swept wings with various aspect ratios and taper 
ratios and with a variety of angle-of-attack or twist 
distributions, including flap and aileron deflections, 
by means of the Weissinger method with eight control 
points on the semispan. Also calculated were aero- 
dynamic influence coefficients which pertain to a cer- 
tain definite set of stations along the span. 



NACA TN 3494 

SOUND PROPAGATION INTO THE SHADOW ZONE 
IN A TEMPERATURE -STRATIFIED ATMOSPHERE 
ABOVE A PLANE BOUNDARY. David C. Pridmore- 
Brown and Uno Ingard, Massachusetts Institute of 
Technology. October 1955. 57p. diagrs., photo. 
(NACA TN 3494) 

A theoretical and ejqierimental study of the sound 
field about a point source over a plane boundary in 
the presence of a vertical temperature gradient has 
been made. Methods are presented for analyzing the 
effects of temperature gradients on the attenuation of 
sound in the shadow zone of a sound field. 



NACA 
RESEARCH 



ABSTRACTS NO. 92 



NACA TN 3495 

FAILURE OF MATERIALS UNDER COMBINED RE- 
PEATED STRESSES WITH SUPERIMPOSED STATIC 
STRESSES. George Sines, University of California 
at Los Angeles. November 1955. 69p. diagrs., 
photos., tabs. (NACA TN 3495) 

Experiments on biaxial alternating stresses and 
simple combinations of static stress with alternating 
stress are reviewed. A general criterion for the 
effect of static stress on the permissible amplitude 
of alternating stress is proposed and compared with 
results of tests performed under more complex 
stress states. Tests were performed to determine 
the effect of static compression on alternating tor- 
sion and the results are compared with the general 
criterion. A modification of Orowan's theory of 
fatigue to include the effect of static stress is pre- 
sented. 



NACA TN 3531 

PILOT'S LOSS OF ORIENTATION IN INVERTED 
SPINS. Stanley H. Scher. October 1955. lOp. 
diagrs., photos. (NACA TN 3531) 

The rising problem of pilot orientation during spins, 
especially during unintentional inverted spins, is 
discussed. The free-spinning-tunnel results and 
reported airplane e.xperiences concerning inverted 
spins and recoveries are reviewed. Information is 
provided regarding the nature of inverted spins, 
optimum control technique for recovery, and some 
of the factors which apparently contribute to the 
pilot's loss of orientation. A spin-simulator rig 
which was recently constructed at the Langley Labo- 
ratory for use in an attempt to understand better the 
problems confronting the pilot of a spinning airplane 
is described. 



NACA TN 3537 

HELICOPTER INSTRUMENT FUGHT AND PRECI- 
SION MANEUVERS AS AFFECTED BY CHANGES IN 
DAMPING IN ROLL, PITCH, AND YAW. James B. 
Whitten, John P. Reeder, and Aimer D. Grim. 
November 1955. 14p. diagrs., photos. 
(NACA TN 3537) 

The damping in roll, pitch, and yaw of a single-rotor 
helicopter was varied by means of electronic compo- 
nents, and these variations were evaluated by per- 
forming instrument approaches and other precision 
maneuvers. Increased damping In roll was found to 
be particularly beneficial, whereas corresponding 
changes In yaw and pitch were less effective. Some 
operational aspects of helicopter Instrument ap- 
proaches are also included In the discussion. 



NACA 
RESEARCH 



ABSTRACTS NO. 92 



NACA TN 3539 

SOME EFFECTS OF SYSTEM NONLINEARITIES IN 
THE PROBLEM OF AIRCRAFT FLUTTER. 
Donald S. Woolston, Harry L. Runyan, and Thomas 
A. Byrdsong. October 1955. 20p. diagrs., tabs. 
(NACA TN 3539) 

This paper presents the results of a preliminary in- 
vestigation of the effect of nonlinear structural 
terms on the flutter of a two-degree-of-freedom 
system. The three types of nonlinearities investi- 
gated were a flat spot, hysteresis, and a cubic 
spring. Calculations were made on an analog com- 
puter. For one case, the flat spot, an experimental 
investigation was also made and good correlation 
with theory was found. 



NACA TN 3542 

ANALYSIS OF STRESSES IN THE PLASTIC RANGE 
AROUND A CIRCULAR HOLE IN A PLATE SUB- 
JECTED TO UNIAXIAL TENSION. Bernard 
Budiansky and Robert J. Vidensek. October 1955. 
39p. diagrs., tabs. (NACA TN 3542) 

An approximate theoretical solution is presented for 
the stresses in the plastic range around a circular 
hole in an infinite sheet subjected to uniaxial tension. 
The solution is based on the simple deformation 
theory of plasticity and is found by application of a 
variational principle in conjunction with the Rayleigh- 
Ritz procedure and the use of a high-speed computing 
machine (SEAC). Numerical results are obtained 
for four different materials, which are characterized 
by four distinct uniaxial stress-strain curves. The 
results for stress concentration factor in the plastic 
range are compared with those obtained from a for- 
mula due to Stowell. 



NACA TN 3544 

COMPARISON BETWEEN THEORETICAL AND EX- 
PERIMENTAL STRESSES IN CIRCULAR SEMI- 
MONOCOQUE CYLINDERS WITH RECTANGULAR 
CUTOUTS. Harvey G. McComb, Jr., and Emmet F. 
Low, Jr. October 1955. 20p. diagrs. 
(NACA TN 3544) 

Comparisons are made between a theory for calcu- 
lating stresses about rectangular cutouts in circular 
cylinders of semimonocoque construction published 
in NACA TN 3200 and previously published NACA 
experimental data. The comparisons include 
stresses in the stringers and shear stresses in the 
center of the shear panels in the neighborhood of the 
cutout. The theory takes into account the bending 
flexibility of the rings in the structure, and this 
factor is found to be important in the calculation of 
stresses about cutouts. In general, when the ring 
flexibility is considered, good agreement is exhibited 
between the calculated and experimental results. 



NACA TN 3550 

MEASUREMENTS OF THE EFFECT OF TRAILING- 
EDGE THICKNESS ON THE ZERO- LIFT DRAG OF 
THIN LOW-ASPECT-RATIO WINGS. John D. 
Morrow. November 1955. lip. diagrs., photo. 
(NACA TN 3550. Supersedes KM L50F26) 

Results of an exploratory free-flight investigation at 
zero lift of several rocket -powered drag-research 
models having 4-percent-thlck wings of taper ratio 
0.423 are presented for a Mach number range of 0.7 
to 1.6. Four wings having trailing edges of different 
thicknesses were tested. The drag of all the models 
was measured and is compared with calculated 
values. 



NACA TN 3557 

A THEORETICAL ANALYSIS OF THE FIELD OF A 
RANDOM NOISE SOURCE ABOVE AN INFINITE 
PLANE. Peter A. Franken, Massachusetts Insti- 
tute of Technology. November 1955. 20p. diagrs. 
(NACA TN 3557) 

The sound field about a random noise source above a 
plane as measured by a receiver with finite band 
width is studied theoretically. For simplicity, only 
the far field is considered. The special case of a 
perfectly reflecting plane is discussed first and the 
analysis is then extended to include the case of a 
plane of arbitrary impedance. 



NACA TN 3567 

STUDY OF SCREECHING COMBUSTION IN A 
6-INCH SIMULATED AFTERBURNER. Perry L. 
Blackshear, Warren D. Rayle, and Leonard K. 
Tower. October 1955. 58p. diagrs., photos., tab. 
(NACA TN 3567) 

The mode of oscillation in a screeching 6-inch- 
diameter simulated afterburner is identified through 
axial, circumferential, and diametric surveys of 
sound amplitude and phase. This mode is found to 
be the first transverse (sloshing) mode in the hot 
gases downstream of the flame-holder. The devel- 
opment of a microphone probe suitable for use in 
screeching combustors is described. This develop- 
ment includes a theoretical and experimental treat- 
ment of the attenuation of high-amplitude sound in 
tubes. 



NACA TN 3572 

AMPLITUDE OF SUPERSONIC DIFFUSER FLOW 
PULSATIONS. William H. Sterbentz and Joseph 
Davids. October 1955. 23p. diagrs. 
(NACA TN 3572. Supersedes RM E52I24) 

A theoretical method for evaluating the stability 
characteristics and the amplitude and frequency of 
pulsation of ram -jet engines without heat addition is 
presented. Theory and experiment show that the 
pulsation amplitude of a high-mass-flow-ratio 
diffuser increases with decreasing mass flow. The 
theoretical trends for changes in amplitude, frequen- 
cy, and mean pressure recovery with changes in 
plenum -chamber volume were experimentally con- 
firmed. For perforated, convergent-divergent -type 
diffusers, theory and experiment show the existence 
of a stability hysteresis loop on the pressure- 
recovery, mass-flow-ratio curve. 



NACA TN 3573 

EFFECT OF EXHAUST-NOZZLE EJECTORS ON 
TURBOJET NOISE GENERATION. Warren J. North 
and Willard D. Coles. October 1955. 26p. diagrs., 
photo. (NACA TN 3573) 

Engine noise levels and jet velocity profiles have 
been obtained with several turbojet exhaust-nozzle 
ejectors. An insignificant reduction in total sound 
power was realized. At subsonic nozzle pressure 
ratios, total sound power from exhaust -nozzle 
ejectors or bypass exit configurations can be calcu- 
lated from primary-jet parameters only. 



NACA TN 3575 

BURNING VELOCITIES OF VARIOUS PREMDCED 
TURBULENT PROPANE FLAMES ON OPEN 
BURNERS. Paul Wagner. October 1955. 32p. 
diagrs., photos., tab. (NACA TN 3575) 

Turbulent burning velocities were measured as a 
function of Reynolds number for open propane 
flames. Flames of propane and oxygen diluted with 
nitrogen, argon, or helium were studied in a variety 
of burners up to a maximum pipe Reynolds number 
of 26,000. The ratio of turbulent to laminar burning 
velocity correlates with the cold-flow Reynolds num- 
ber for systems of a given diluent. This ratio also 
correlates with a Reynolds number calculated from 
values of the turbulent intensity measured at the 
burner exit. 



NACA 
RESEARCH 



ABSTRACTS NO. 92 



BRITISH REPORTS 



N-40036* 

Royal Aircraft Establishment (Gt. Brit.) 
LOW SPEED WIND TUNNEL CALIBRATION OF A 
Mk. 9A PITOT -STATIC HEAD. J. E. Nethaway. 
March 1955. 8p. diagrs., tab. (RAE Tech. Note 
Aero 2364) 

This note describes the calibration of a Mark 9A 
pitot-static head in the No. 2 11-1/2 foot wind 
tunnel. The results show the variation of pitot and 
static-pressure error coefficients with incidence, at 
constant tunnel speed. 



N-40039* 

Royal Aircraft Establishment (Gt. Brit.) 
THE DETERMINATION OF THE RELATIVE REAC- 
TIVITIES OF BIFUNCTIONAL MONOMERS. PART 
I - THE NATURE OF THE PROBLEM. G. M. 
Bristow. July 1955. 6p. tab. (RAE Tech. Note 
Chem. 1256) 

An account is given of the use of bifunctional mono- 
mers in the production of cross-linked polymers, 
and of some methods by which the relative reactivi- 
ties of the two unsaturated groups can be determined. 



N-40042* 

Royal Aircraft Establishment (Gt. Brit.) 
A LOW-PRESSURE MICRO-ANALYTICAL METHOD 
FOR DETERMINING OXYGEN, HYDROGEN AND 
NITROGEN IN METALS. H. C. Davis and J. A. 
Gray. March 1955. 14p. diagrs., tabs. (RAE 
Met. 86) 

A low-pressure microanalytical method for the 
determination of oxygen, hydrogen, and nitrogen in 
metals is described. Gases evolved by vacuum- 
fusion are analyzed physically. The results obtained 
for oxygen and hydrogen are closely reproducible, 
but those for nitrogen are less so. 



NACA 

RESEARCH ABSTRACTS NO. 92 



N-40044* 



Royal Aircraft Establishment (Gt. Brit.) 
THE THEORETICAL WAVE DRAG OF OPEN NOSE 
AXISYMMETRICAL FOREBODIES WITH VARYING 
FINENESS RATIO, AREA RATIO AND NOSE ANGLE. 
J. H. Willis and D. G. Randall. February 1955. 
33p. diagrs., tab. (RAE Tech. Note Aero 2360) 

Existing results for the wave drag of open-nose 
axisymmetrical forebodies are for bodies whose 
profiles are straight lines or parabolic arcs. These 
results are here extended to a family of profiles 
which includes the straight line and the parabolic arc 
as special cases. Slender body theory is employed 
throughout. 



UNPUBLISHED PAPERS 



N-40045' 

Royal Aircraft Establishment (Gt. Brit.) 
TESTS ON HEAT AND CORROSION RESISTING 
COATINGS FOR MAGNESIUM ALLOYS. J. Mackay 
and H. G. Cole. June 1955. 9p. tabs. (RAE 
Tech. Note Chem. 1254) 

In tests on painted magnesium alloy specimens 
heated to 200° C at three-monthly intervals during 
an intermittent seawater spray test, the best com- 
binations of protective efficiency with hardness and 
flexibility were given by an epoxy base scheme and 
by a standard stoving scheme to D. T. D. 235. 
Other epoxy schemes gave good protection but were 
relatively inflexible. A standard air-drying scheme 
toD.T.D. 260A gave very good protection. Butyl 
titanate paints and a silicone paint gave poor pro- 
tection. A zinc chrome sealed anodic treatment 
gave better results as a protective pretreatment 
than chromate bath iii of D.T.D. 911A. Butyl titan- 
ate paints broke down on pure aluminum exposed to 
the same test. 



N-26685* 

EXPERIMENTAL DETERMINATION OF THE EF- 
FECTIVE WIDTH OF FLAT PLATES IN THE ELAS- 
TIC AND PLASTIC RANGE. (De experimentele 
bepaling van de meedragende breedte van vlakke 
platen in het elastische en het plastische gebied). 
J. F. Besseling. September 1955. lOOp. diagrs., 
photos., tab. (Trans, from Nationaal Luchtvaart- 
laboratorium, Amsterdam, S.414, February 1953) 

Influence of boundary conditions and transverse 
stiffeners on the effective width of a plate are dis- 
cussed. The testing apparatus, measuring equip- 
ment, and test program are reviewed and test re- 
sults are presented for aluminum alloy 24ST. 



N-39572 

HEAT-RESISTANT SINTER MATERIALS. 
(Hochwarmfeste Sinterwerkstoffe). F. Benesovsky. 
(1955). 6p. diagrs., photo., tab. (Trans, from 
Werkstoffe und Korrosion, Jour., no. 8/9, Aug. - 
Sept. 1954, p. 288-290) 

Ceramal materials, produced by powder metallurgy 
from a titanium carbide base, are suggested for use 
in turbine and rocket drives. Other ceramal mate- 
rials suggested for high temperature use are alloys 
with boride, silicide, and oxide bases. Titanium 
boride and zirconium boride show high melting 
points, are generally brittle, and have poor thermal 
shock properties. It is believed that metal additions 
may improve these properties. 



MISCELLANEOUS 



NACA Rept. 1135 

Errata No. 2 on "EQUATIONS, TABLES, AND 
CHARTS FOR COMPRESSIBLE FLOW. " Ames 
Research Staff. 1953. 



NACA Rept. 1175 

Errata on "EFFECT OF VARIABLE VISCOSITY AND 
THERMAL CONDUCTIVITY ON HIGH-SPEED SLIP 
FLOW BETWEEN CONCENTRIC CYLINDERS. ' 
T. C. Lin and R. E. Street. 1954. 



NACA TN 3454 

Errata on "EFFECT OF A DISCONTINUITY ON 
TURBULENT BOUNDARY-LAYER-THICKNESS 
PARAMETERS WITH APPLICATION TO SHOCK- 
INDUCED SEPARATION. " Eli Reshotko and 
Maurice Tucker. May 1955. 



NACA TN 3483 

Errata on "AN ANALYSIS OF ACCELERATION. 
AIRSPEED, AND GUST-VELOCITY DATA FROM A 
FOUR-ENGINE TRANSPORT AIRPLANE IN OPERA- 
TIONS ON AN EASTERN UNITED STATES ROUTE. * 
Thomas L. Coleman and Mary W. Fetner. 
September 1955. 



NACA TN 3514 

Errata on "RESPONSE OF HOMOGENEOUS AND 
TWO-MATERIAL LAMINATED CYLINDERS TO 
SINUSOIDAL ENVIRONMENTAL TEMPERATURE 
CHANGE, WITH APPLICATIONS TO HOT-WIRE 
ANEMOMETRY AND THERMOCOUPLE PYROME- 
TRY. " Herman H. Lowell and Norman Patton. 
September 1955. 



N-37765* 

SOME ELEVATED TEMPERATURE STRUCTURAL 
PROBLEMS OF HIGH-SPEED AIRCRAFT. 
Richard R. Heldenfels. (Presented to SAE 
Golden Anniversary National Aeronautic Meeting, 
Los Angeles, California, October 11-15, 1955) 29p. 
diagrs. 

This paper contains a discussion of the problen.s of 
structural design, such as creep, thermal uuckliiig, 
thermal stresses, and stiffness reduction, as related 
to the aerodynamic heating of high-speed aircraft. 



DECLASSIFIED NACA REPORTS 



THE FOLLOWING REPORTS HAVE BEEN 
DECLASSIFIED FROM CONFIDENTIAL, 10/14/55 



NACA RM A7J02 

THE HIGH-SPEED AERODYNAMIC EFFECTS OF 
MODIFICATIONS TO THE WING AND WING- 
FUSELAGE INTERSECTION OF AN AIRPLANE 
MODEL WITH THE WING SWEPT BACK 35°. Lee 
E. Boddy and Charles P. Morrill, Jr. February 1£ 
1948. 34p. diagrs.. photos. (NACA RM A7J02) 

Wind-tunnel tests at high subsonic Mach numbers 
were conducted on an airplane model having swept - 
back wings. Attempts were made to reduce the 
interference at the plane of symmetry of the swept- 
back wing and thus increase its divergence Mach 
number. Also, tests were made with the wing 
trailing-edge angle decreased, in an effort to elimi- 
nate the reversal of aileron hinge moment and wing 
pitching moment suffered by the true-contour wing 
at high Mach numbers. 



NACA 
RESEARCH 



ABSTRACTS NO. 92 



NACA RM A7J03 

THE AERODYNAMIC EFFECTS OF ROCKETS AND 
FUEL TANKS MOUNTED UNDER THE SWEPT- 
BACK WING OF AN AIRPLANE MODEL. Lee E. 
Boddy and Charles P. Morrill, Jr. April 23, 1948. 
19p. diagrs. (NACA RM A7J03) 

The effects of externally mounted rockets and fuel 
tanks on the aerodynamic characteristics of an air- 
plane model with sweptback wings are presented in 
this report. Wind-tunnel tests were made at high 
subsonic Mach numbers to determine the effect of 
the external equipment on the drag, the longitudinal 
stability and control, and the lateral control of the 
model. 



NACA RM A50J26 

EXPERIMENTAL DAMPING IN PITCH OF 45° TRI- 
ANGULAR WINGS. Murray Tobak, David E. Reese, 
Jr., and Benjamin H. Beam. December 1, 1950. 
63p. diagrs., photo. (NACA RM A50J26) 

Results are presented of a wind-tunnel investigation 
of the variation of the damping-ln-pitch parameter 
•^m + C_ with Mach number and axis of rotation 

position for two triangular wings having leading 
edges swept back 45°, with and without a body. 
Tests were conducted at subsonic speeds over a 
Mach number range of 0.23 to 0.94 and at supersonic 
speeds from 1.15 to 1.70. The measured damping 
coefficients are compared with theoretical results at 
both subsonic and supersonic speeds. 



NACA RM A53J02 

PRELIMINARY RESULTS OF AN INVESTIGATION 
OF THE EFFECTS OF SPINNER SHAPE ON THE 
CHARACTERISTICS OF AN NACA D-TYPE COWL 
BEHIND A THREE-BLADE PROPELLER, IN- 
CLUDING THE CHARACTERISTICS OF THE PRO- 
PELLER AT NEGATIVE THRUST. Robert M. 
Reynolds. November 1953. 15p. diagrs., photo., 
tab. (NACA RMA53J02) 

Preliminary results of measurements of the ram- 
recovery ratio at the inlet of an NACA D-type cowl 
behind an operating propeller in combination with a 
1 -series and a modified-conical spinner, maximum 
efficiency of the propeller with the 1 -series spinner 
and the spinner-cowling combinations, and the nega- 
tive thrust characteristics of the propeller at low 
speeds are summarized. Tests were conducted at 
Mach numbers from 0.2 to 0.8, for propeller blade 
angles from 33° to 63°, and for various inlet- 
velocity and advance ratios. Negative-thrust char- 
acteristics of the propeller were measured at a Mach 
number of 0.15 for blade angles from 25° to -20°. 
All tests were made with the model at an angle of 
attack of 0° and at a Reynolds number of 1.0 million 
per foot. 



NACA 
RESEARCH 



ABSTRACTS NO. 92 



NACA RM A53J07 

THE EFFECTS OF HORIZONTAL-TAIL HEIGHT 
AND A PARTIAL-SPAN LEADING-EDGE EXTEN- 
SION ON THE STATIC LONGITUDINAL STABILITY 
OF A WING-FUSELAGE-TAIL COMBINATION 
HAVING A SWEPTBACK WING. Angelo Bandettini 
and Ralph Selan. March 1954. 54p. diagrs., 
photos., 2 tabs. (NACA RM A53J07) 

Test results are presented to show the effects of 
horizontal-tail height (22- and 8-percent semispan 
above wing chord plane extended) on the static longi- 
tudinal stability of a model having a wing with 35° 
sweepback, an aspect ratio 4.5, and a taper ratio 0.5. 
The model was also modified by a wing-leading-edge 
chord extension and tested with the tail in the low 
position. Tests were conducted at various Mach 
numbers up to 0.92 at a Reynolds number of 2,000,000 
and at a Mach number of 0.20 at a Reynolds number 
of 11,000,000. The results of airstream surveys in 
the region of the tail are also presented. 



NACA RM E50J24 

DYNAMIC INVESTIGATION OF TURBINE - 
PROPELLER ENGINE UNDER ALTITUDE CONDI- 
TIONS. Richard P. Krebs, Seymour C. Himmel, 
Darnold Blivas, and Harold Shames. December 6, 
1950. 55p diagrs., photo. (NACA RM E50J24) 

An altitude-wind-tunnel investigation of the dynamics 
of a turbine -propeller engine employing the 
frequency-response technique was conducted over a 
range of pressure altitudes from 10,000 to 30,000 
feet. The dynamic responses generalized for pres- 
sure altitudes over the range of frequencies investi- 
gated. The generalized time constants were found to 
be approximately 1.0 second for the engine-propeller 
combination, 0.36 second for the propeller alone, and 
2.4 seconds for the engine alone. These values were 
in good agreement with those predicted from steady - 
state-performance data. 



NACA RM E51J11 

IGNITION-DELAY CHARACTERISTICS IN MODIFIED 
OPEN-CUP APPARATUS OF SEVERAL FUELS 
WITH NITRIC ACID OXIDANTS WITHIN TEMPERA- 
TURE RANGE 70° TO -105° F. Riley O. Miller. 
December 1951. 30p. diagrs., 4 tabs. 
(NACA RM E51J11) 

Fluid properties and low-temperature ignition delays 
were obtained for approximately 90 fuel-oxidant com- 
binations. A red fuming nitric acid containing ap- 
proximately 3 percent water and 19 percent nitrogen 
tetroxide froze at approximately -87° F and ignited 
several low-viscosity fuel blends of aromatic amines 
in triethylamine at -76° F and lower. With this acid, 
the following average ignition delays were obtained 
with a blend of 30 percent o-toluidine in triethyla- 
mine: 



Temperature, OF 
Delay, milliseconds 


70 
19 


-40 
24 


-76 
38 


-87 
61 


-105 
210 



NACA RM E52J27 

PRELIMINARY INVESTIGATION OF A PERFO- 
RATED AXIALLY SYMMETRIC NOZZLE FOR 
VARYING NOZZLE PRESSURE RATIOS. Eli 
Reshotko. January 1953. 43p. diagrs., photo., 
2 tabs. (NACA RM E52J27) 

The performance characteristics of a perforated 
axially symmetric convergent -divergent nozzle were 
investigated in an attempt to achieve improved 
convergent-divergent nozzle thrust performance at 
below design pressure ratios. The purpose of the 
perforations was to allow inflow of air into the over- 
expanded portion of the nozzle, thus advancing sepa- 
ration of the flow. The flow through the perfora- 
tions was found to advance separation only when the 
perforations were liberally placed over the entire 
divergent portion of the nozzle. A local concentra- 
tion of perforations caused separation only in the 
local region of perforation. The use of low energy 
atmospheric bleed air reduced thrust losses by as 
much as 50 percent at appreciably overe.xpanded 
operation. For underexpanded flow, air flowing out 
through the perforations caused significant thrust 
loss. With shrouding to prevent this outbleed, 
thrusts 5 to 10 percent less than those of the un- 
perforated nozzle were obtained. The use of high 
energy bleed was unsatisfactory since the inlet 
momentum penalty of the bleed air was in many 
cases greater than the additional thrust obtained. 



NACA RM L8I29 

PRELIMINARY RESULTS OF NACA TRANSONIC 
FLIGHTS OF THE XS-1 AIRPLANE WITH 10- 
PERCENT-THICK WING AND 8-PERCENT-THICK 
HORIZONTAL TAIL. Hubert M. Drake, Harold R. 
Goodman, and Herbert H. Hoover. October 13, 1948. 
18p. diagrs., photos. (NACA RM L8I29) 

Contains results of exploratory flights at altitudes of 
about 40,000 feet to a maximum Mach number of 
1.06. Data are presented showing the longitudinal 
trim changes, elevator effectiveness in producing 
acceleration, and rudder effectiveness as a Junction 
of Mach number. Data on lateral oscillations are 
also presented. 



NACA RM L8J12 

HIGH-SPEED WIND-TUNNEL INVESTIGATION OF 
A SWEPTBACK WING WITH AN ADDED TRIANGU- 
LAR AREA AT THE CENTER. Beverly Z. Henry, 
Jr. January 14, 1949. 24p. diagrs., tabs. 
(NACA RM L8J12) 

Results are presented of an investigation in the 
Langley 8-foot high-speed tunnel of two sweptback 
wings of different plan form. The purpose of the 
investigation was to determine the effects of the 
addition of a triangular area to the inboard section 
of a conventional sweptback wing in order to produce 
a wing employing two stages of sweepback. Lift, 
drag, and pitching-moment characteristics are pre- 
sented to Illustrate these effects for a Mach number 
range of 0.40 through 0.935. 



NACA RM L8K23 

FREE-FLIGHT INVESTIGATION AT TRANSONIC 
AND SUPERSONIC SPEEDS OF THE ROLLING EF- 
FECTIVENESS OF SEVERAL AILERON CONFIGU- 
RATIONS ON A TAPERED WING HAVING 42.7° 
SWEEPBACK. Carl A. Sandahl. January 11, 1949. 
23p. diagrs., photos., tab. (NACA RM L8K23) 

An investigation was made of several aileron modi- 
fications in conjunction with a tapered, sweptback 
wing having circular-arc airfoil sections of rela- 
tively large thickness ratio. The modified ailerons 
eliminated the reversal of rolling effectiveness at 
transonic speeds obtained with the true-contour ai- 
lerons at small deflections. 



NACA RM L9F07 

PRELIMINARY THEORETICAL AND FLIGHT IN- 
VESTIGATION OF THE LATERAL OSCILLATION 
OF THE X-1 AIRPLANE. Hubert M. Drake and 
Helen L. Wall. July 19, 1949. 24p. diagrs., photo., 
tab. (NACA RM L9F07) 

A small -amplitude, undamped, lateral oscillation has 
been encountered in flight tests of the X-1 airplane. 
The oscillation occurs in subsonic and supersonic 
flight, in maneuvers, and power on and off. The 
calculations indicate that a change, in the positive 
direction, of the inclination of the principal axis with 
respect to the flight path should have a considerable 
stabilizing effect. 



NACA RM L9G19a 

MEASUREMENTS OF AILERON EFFECTIVENESS 
OF THE BELL X-1 AIRPLANE AT MACH NUMBERS 
BETWEEN 0.9 AND 1.06. Hubert M. Drake. 
August 4, 1949. 5p. diagrs. (NACA RM L9G19a) 

Presents results of flight measurements of aileron 
effectiveness of the X-1 airplane up to a Mach num- 
ber of 0.94. The data indicate a 75 percent loss of 
aileron effectiveness between M = 0.82 and 
M = 0.94. 



NACA RM L50G20 

ELEVATOR-STABILIZER EFFECTIVENESS AND 
TRIM OF THE X-1 AIRPLANE TO A MACH NUM- 
BER OF 1.06. Hubert M. Drake and John R. Garden. 
November 1, 1950. 12p. diagrs. 
(NACA RM L50G20) 

The relative elevator -stabilizer effectiveness of the 
X-1 has been determined to decrease from a value of 
0.25 at a Mach number of 0.78 to a value of 0.05 at a 
Mach number of 1.0. At supersonic speeds the ef- 
fectiveness increases. The variation between the 
trim curves at various stabilizer settings is caused 
by the variation in effectiveness and the fact that the 
effectiveness is nonlinear at Mach numbers between 
0.94 and 0.97. It was found that, with the elevator 
fixed at zero, only about 0.5° of stabilizer movement 
would be required to trim through the Mach number 
range from 0.78 to 1.02. 



NACA 
RESEARCH 



ABSTRACTS NO. 92 



NACA RM L50J25 



A TRANSONIC -WING INVESTIGATION IN THE 
LANGLEY 8-FOOT HIGH-SPEED TUNNEL AT HIGH 
SUBSONIC MACH NUMBERS AND AT A MACH NUM- 
BER OF 1.2. WING-FUSELAGE CONFIGURATION 
HAVING A WING OF 60° SWEEPBACK, ASPECT 
RATIO 4, TAPER RATIO 0.6, AND NACA 65A006 
AIRFOIL SECTION. Raymond B. Wood and Frank F. 
Fleming. January 24, 1951. 43p. diagrs., photo. 
(NACA RM L50J25) 

An investigation was conducted in the Langley 8-foot 
high-speed tunnel of the aerodynamic characteristics 
of a wing swept back 60° at the quarter chord, with 
aspect ratio 4, taper ratio 0.6, and an NACA 65A006 
airfoil section. The tests were conducted through a 
Mach number range from 0.6 to 0.96 and at a Mach 
number of 1.2. Data are presented for a wing 
fuselage and for a wing with wing-fuselage interfer- 
ence. Wake-survey-study results and the measure- 
ments of the angle of downwash for a probable tail 
location, approximately 38 percent of the wing semi- 
span above the wing-chord plane, are included. 



NACA RM L51D17 

AN INVESTIGATION OF FOUR WINGS OF SQUARE 
PLAN FORM AT A MACH NUMBER OF 6.86 IN THE 
LANGLEY 11-INCH HYPERSONIC TUNNEL. 
Charles H. McLellan, Mitchel H. Bertram, and 
John A. Moore. June 1951. 47p. diagrs., photos. 
(NACA RM L51D17) 

The results of tests of four wings at a Mach number 
of 6.86 in the Langley 11-inch hypersonic tunnel are 
presented. The wings tested had 4-inch square plan 
forms with 5-percent-thick diamond, half-diamond, 
wedge, and half-circular-arc sections. The bound- 
ary layer has been found to have a large effect on the 
wing pressure distributions. Reasonable agreement 
was indicated between the aerodynamic coefficients 
from experimental pressure data and inviscid theory. 
Total drag measurements showed good agreement 
with the theory at low angles of attack when the ef- 
fects of surface friction were included. At the 
higher angles of attack, both lift coefficient and drag 
coefficient were found to be slightly below the values 
predicted by the two-dimensional theory. 



NACA RM L51J10 

INVESTIGATION BY THE TRANSONIC -BUMP 
METHOD OF A 35° SWEPTBACK SEMISPAN MODEL 
EQUIPPED WITH A FLAP OPERATED BY A SERIES 
OF SERVOVANES LOCATED AHEAD OF AND 
GEARED TO THE FLAP. William H. Phillips and 
Robert F. Thompson. December 1951. 39p. 
diagrs., photo. (NACA RM L51J10) 

Lift. drag, pitching-moment. rolling-moment, and 
yawing-moment data in the Mach number range from 
0.6 to 1.0 obtained from wind-tunnel tests of a low- 
aspect-ratio sweptback airfoil model with a servo- 
vane control are presented. The control utilizes 
the drag force and spoiler action of a set of vanes to 
deflect a flap-type control. Comparison of lift in- 
crement and center-of-pressure location is made 
with previously published data from tests of a con- 
ventional flap-type control. 



NACA 
RESEARCH 



ABSTRACTS NO. 92 



UNIVERSITY OF FLORIDA 




1262 08153 288 8 



NACA RM L51J30 

SUMMARY OF FLUTTER EXPERIENCES AS A 
GUIDE TO THE PRELIMINARY DESIGN OF LIFTING 
SURFACES ON MISSILES. Dennis J. Martin. 
November 1951. 16p. diagrs. (NACA RM L51J30) 

This report presents a limited review of some ex- 
periences in flight testing of missiles and of wing 
flutter investigations that may be of interest in mis- 
sile design. Several types of flutter which may be of 
concern in missile studies are briefly described. 
Crude criteria are presented for two of the most 
common types of flutter to permit a rapid estimate 
to be made of the probability of the occurrence of 
flutter. Many of the details of the flutter picture 
have been omitted, and only the broader elements 
have been retained so as to give the designer an 
overall view of the subject. 



NACA RM L52D01 

A STUDY OF THE FLOW OVER A 45° SWEPTBACK 
WING-FUSELAGE COMBINATION AT TRANSONIC 
MACK NUMBERS. Richard T. Whitcomb and 
Thomas C. Kelly. June 1952. 60p. diagrs., 
photos. (NACA RM L52D01) 

Pressure distributions, tuft patterns, and schlieren 
surveys have been obtained for a 45° sweptback 
wing-fuselage combination in the Langley 8-foot 
transonic tunnel at transonic Mach numbers to 1.11 
and angles of attack to 20°. The results provide an 
indication at transonic Mach numbers of the nature 
of the formation of shock waves on the wing and 
fuselage, wing-fuselage interference, and the de- 
velopment of separation and the separation vortex. 



NACA RM L52J21a 

INVESTIGATIONS AT SUPERSONIC SPEEDS OF THE 
BASE PRESSURE ON BODIES OF REVOLUTION 
WITH AND WITHOUT SWEPTBACK STABILIZING 
FINS. Eugene S. Love and Robert M. O'Donnell. 
December 1952. 66p. diagrs., photos. 
(NACA RM L52J21a) 

Results are presented from an investigation at Mach 
numbers of 1.62, 1.93, and 2.41 of the variation 
with Reynolds number of the base pressure on bodies 
of revolution at zero lift, with and without sweptback 
stabilizing fins. Included are the effects of varying 
nose and base shapes and cutoff length, the effects 
of the presence of sting supports of varying diam- 
eter, and the effects of disturbances entering the 
wake. The overall Reynolds number range was ap- 
proximately from 1 X 10^ to 10 x 10^. 



NACA RM L52J23a 

INVESTIGATION OF THE AERODYNAMIC CHARAC- 
TERISTICS OF THE NACA RM-10 MISSILE (WITH 
FINS) AT A MACH NUMBER OF 1.62 IN THE 
LANGLEY 9-INCH SUPERSONIC TUNNEL. Donald 
E. Coletti. December 1952. 21p. diagrs. 
(NACA RM L52J23a) 

An investigation was made of a 0.050-scale model of 
the RM-10 missile at a Mach number of 1.62 and a 



Reynolds number of 2.66 x 10^. Measurements were 
made of lift, drag, and pitching moment over an 
angle-of-attack range of t5°. The effects of the 
ratio of sting-shield diameter to base diameter were 
also investigated. Comparisons are made with re- 
sults of tests in other facilities at widely different 
Reynolds numbers. 



NACA RM L52K06 

PRESSURE DISTRIBUTION AND PRESSURE DRAG 
FOR A HEMISPHERICAL NOSE AT MACH NUM- 
BERS 2.05. 2.54, AND 3.04. Leo T. Chauvin. 
December 1952. 14p. diagrs., photos. 
(NACA RM L52K06) 

An experimental investigation of the pressure dis- 
tributions on a hemispherical nose 3.98 inches in 
diameter, mounted on a cylindrical support, has 
been made at Mach numbers of 2.05, 2.54, and 3.04 
and for Reynolds numbers of 4.44 x 10^, 4.57 x 10^, 
and 4.16 x 10^, respectively. The Reynolds number 
was based on body diameter and free-stream condi- 
tions. Pressure-drag coefficients were calculated 
and good agreement was obtained between these tests 
and other investigations. 



NACA RM L53E04 

AILERON AND ELEVATOR HINGE MOMENTS OF 
THE BELL X-1 AIRPLANE MEASURED IN TRAN- 
SONIC FLIGHT. Hubert M. Drake and John B. 
McKay. June 1953. 27p. diagrs. 
(NACA RM L53E04) 

Hinge moments have been measured on the aileron 
and elevator of the Bell X-1 airplane having the 10- 
percent-thick wing and 8-percent-thick tail. The 
aileron measurements were made by means of 
strain gages and pressure distributions while the 
elevator measurements were made by means of the 
wheel-force strain gages. The elevator hinge- 
moment characteristics were determined to a Mach 
number of 1. 18 and the aileron hinge moments to a 
Mach number of 1. 13. 



NACA RM L53F08 

FLIGHT MEASUREMENTS OF LIFT AND DRAG 
FOR THE BELL X-1 RESEARCH AIRPLANE HAV- 
ING A 10-PERCENT-THICK WING. Edwin J. 
Saltzman. September 1953. 37p. diagrs., tab. 
(NACA RM L53F08) 

Lift and drag results have been obtained from power- 
off flight tests of the Bell X-1 (lO-percent-thick 
wing) airplane for Mach numbers 0.68 to 1.01. Com- 
parisons of drag are made with 8-percent-thick-wing 
flight tests and 10-percent-thick-wing wind-tunnel 
results. 



10 



NACA RM L53I09a 

LOW-SPEED WIND-TUNNEL INVESTIGATION OF A 
JET CONTROL ON A 35° SWEPT WING. John G. 
Lowry and Thomas R. Turner. October 1953. 9p. 
diagrs. (NACA RM L53I09a) 

A low-speed wind-tunnel investigation was made of a 
jet control that obtains its effectiveness from both 
the jet reaction and from the change in circulation 
around the wing due to the jet's acting as a spoiler. 
The jet control was investigated as an aileron on a 
35° sweptback wing of aspect ratio 4.76. The in- 
vestigation was of exploratory nature and was limited 
to the case where the jet was supplied with air at 
stagnation pressure. The results indicated that such 
a jet could be used as an emergency control. 



NACA RM L53J01a 

WIND-TUNNEL INVESTIGATION OF THE EFFECTS 
OF STEADY ROLLING ON THE AERODYNAMIC 
LOADING CHARACTERISTICS OF A 45° SWEPT- 
BACK WING AT HIGH SUBSONIC SPEEDS. James 
W. Wiggins and Richard E. Kuhn. November 1953. 
22p. diagrs., photos. (NACA RM L53J01a) 

The aerodynamic loading characteristics of a 45 
sweptback wing of aspect ratio 4 in combination with 
a fuselage during steady roll are presented. The 
tests covered Mach numbers of 0.70, 0.85, and 0.91, 
and angles of attack up to 13°. The effects of fences 
at a Mach number of 0.85 and a comparison of 
measured and calculated load distribution are in- 
cluded. 



NACA RM L53J09a 

WIND-TUNNEL INVESTIGATION AT LOW SPEED 
OF THE EFFECT OF VARYING THE RATIO OF 
BODY DIAMETER TO WING SPAN FROM 0.1 TO 0.8 
ON THE AERODYNAMIC CHARACTERISTICS IN 
PITCH OF A 45° SWEPTBACK-WING— BODY 
COMBINATION. Harold S. Johnson. November 
1953. 32p. diagrs., photo., tab. 
(NACA RM L53J09a) 

Low-speed lift, drag, and pitching-moment data 
were obtained for a family of bodies and wing-body 
combinations to determine the effect of varying the 
ratio of body diameter to wing span from 0.1 to 0.8. 
The bodies had ogival noses and cylindrical after- 
bodies. The untapered 6-percent-thick wings had 
aspect ratios of 3 and 45° of sweepback. The lift 
data of the bodies alone and the wing-body combina- 
tions are compared with several existing theories. 



NACA RM L53J19 

AN EXPERIMENTAL AND THEORETICAL INVES- 
TIGATION AT HIGH SUBSONIC SPEEDS OF THE 
EFFECTS OF HORIZONTAL-TAIL HEIGHT ON THE 
AERODYNAMIC CHARACTERISTICS IN SIDESLIP 
OF AN UNSWEPT, UNTAPERED TAIL ASSEMBLY. 
Harleth G. Wiley and Donald R. Riley. December 
1953. 71p. diagrs., tab. (NACA RM L53J19) 

This paper presents the effects at high subsonic 
speeds of horizontal-tail height on the aerodynamic 
characteristics in sideslip at 0° angle of attack of an 



NACA 
RESEARCH 



ABSTRACTS NO. 92 



unswept, untapered empennage. Configurations in- 
vestigated included the fuselage alone, fuselage plus 
vertical tail, fuselage plus horizontal tail, and the 
fuselage plus vertical tail with the horizontal tail 
located at 0, 26, 59, and 100 percent vertical- 
surface span. Tests were made at 0° angle of attack 
through a sideslip range of -2° to 20° over a Mach 
number range of 0.50 to 0.94. 



NACA RM L53J29 

WIND-TUNNEL INVESTIGATION AT HIGH AND LOW 
SUBSONIC MACH NUMBERS OF TWO UNSWEPT 
WINGS HAVING NACA 2-006 AND NACA 65A006 
AIRFOIL SECTIONS. Stanley F. Racisz. 
December 1953. 40p. diagrs., photo., tab. 
(NACA RM L53J29) 

An investigation has been made of two unswept wings 
with aspect ratios of 4 and taper ratios of 0.2. One 
wing had airfoil sections designed for high maximum 
lift at low speeds (NACA 2-006), and the other wing 
had NACA 65A006 airfoil sections. Each wing was 
mounted on a slender body of revolution. The lift, 
drag, and pitching-moment characteristics were 
determined at Reynolds numbers from 1 x 10^ to 
7.5 X 10^ for Mach numbers below 0.2 for the wings 
with and without split flaps and for the wings with and 
without leading-edge roughness. The character- 
istics of the plain wings were also determined for 
several values of Reynolds number at Mach numbers 
up to about 0.92. Gains obtainable by the use of the 
NACA 2-006 airfoil section are evident for Mach 
numbers up to 0.65 from the comparisons of the re- 
sults for the two wings. 



NACA RM L53L08a 

EXPERIMENTAL CONVECTIVE HEAT TRANSFER 
TO A 4-INCH AND 6-INCH HEMISPHERE AT MACH 
NUMBERS FROM 1.62 TO 3.04. Leo T. Chauvin 
and Joseph P. Maloney. February 1954. ISp. 
diagrs., photos. (NACA RM L53L08a) 

Equilibrium temperatures and heat-transfer coeffi- 
cients for a hemispherical nose have been measured 
for Mach numbers from 1.62 to 3.04. Heat transfer 
to the surface of the hemisphere was presented as 
Stanton number against Reynolds number for various 
surface heating conditions. Heat transfer at the 
stagnation point has been measured and correlated 
with theory. Transition from a laminar to a turbu- 
lent boundary layer was obtained at Reynolds num- 
bers of approximately 1 x 10^ corresponding to a 
region on the body located between 45° and 60° from 
the stagnation point. 



NACA RM L53L15 

INVESTIGATION OF A PULSE-JET-POWERED 
HELICOPTER ROTOR ON THE LANGLEY HELI- 
COPTER TEST TOWER. Edward J. Radin and 
Paul J. Carpenter. February 1954. 23p. diagrs., 
photos. (NACA RM L53L15) 

A helicopter rotor powered by tip-located pulse-jet 
engines has been investigated on the Langley heli- 
copter test tower to determine its basic hovering 
characteristics as well as the power-off drag and 
propulsive characteristics of the engines. The noise 
intensity in the vicinity of the pulse-jet engines was 
also determined. 

NACA - Langley Field, Va.